1. Field of the Invention
This invention relates to the field of mechanical design, and in particular to the design of a deployment device that is configurable to provide a substantially constant deployment velocity. Such a device is particularly well suited for deployment of equipment associated with a spacecraft.
2. Description of Related Art
Hinges and other devices are commonly used to facilitate the deployment of equipment from a ‘stored’ state to a ‘deployed’ state. In a spacecraft system, for example, subsystems are stored in a compact form to facilitate stowage of the spacecraft in a launch vehicle. When the spacecraft is deployed to its operational orbit, or operational trajectory, the subsystems that support the mission of the spacecraft are deployed to their operational form. Antenna systems are deployed to provide a desired field of view; solar panels are deployed to provide a surface area that is sufficient to collect the energy required to power the equipment on the space-craft; sensing devices are deployed at a relatively large distance from the space-craft to collect data that is substantially independent of the effects of the space-craft; and so on. Deployment systems are commonly used in other fields, but the field of aerospace places particularly stringent demands on the reliability and efficiency of space-craft deployment systems, and thus this invention is presented using the paradigm of a hinged deployment device that is particularly well suited for space-craft design.
Deployment systems generally face a set of conflicting requirements that are generally related to the ‘starting’ and ‘stopping’ of the deployment. The deployment must be reliably accomplished, and thus more force than is minimally required is typically used to effect the deployment, to provide a sufficient ‘reliability margin’. Typically, this extra force is transferred to the mass in the form of excess kinetic energy. This excess kinetic energy must be dissipated at the end of deployment. A sudden stop of a rotating mass about a hinge introduces a substantial shock stress on the hinge and on the structure that is used to stop the mass. A sudden stop of an expanding telescope introduces a substantial shock stress on the limits used at each telescoping element.
Shock absorbers are commonly used to dampen the effects associated with stopping a moving mass, by stopping the mass gradually. An overdamped deployment system will generally cause a lack of full deployment and an underdamped deployment system does not fully abate the aforementioned shock effects. Generally, the costs of a failed deployment are significant, and most systems are purposely designed to be underdamped, and the design of the deployment system includes considerations for withstanding substantial shock effects.
An ideal deployment system is one that is able to apply a large amount of force, as required, to move the mass under a variety of non-ideal conditions, yet limit the terminal velocity of the mass at or below a given rate to minimize the amount of kinetic energy that must be dissipated to stop the motion of the mass.
An electric motor can be configured to approximate the characteristics of an ideal deployment system, via a regulated control of the energy/current that is applied to the motor to provide the appropriate torque and velocity. Such a solution, however, is often economically infeasible. In a spacecraft environment, the added mass and complexity of a motor also introduces other design considerations and tradeoffs. In the case of a solar-panel deployment system, for example, these additional considerations include the need to provide power to the motor before the solar panels are deployed.
Most commonly, tensioned springs are used in a deployment system. A tensioned coil spring about the axle of a hinge, for example, provides a simple and reliable means of supplying the force required to rotate a mass about the hinge. A compressed linear coil spring applied to a piston provides a simple and reliable means of supplying the force required to move a mass in a linear direction. The spring is typically designed to provide more than enough force to overcome a worst-case scenario of adverse conditions. As such, a substantial amount of dampening is required to minimize or reduce shock effects, as discussed above.
Commonly, dampening devices are fluidic or pneumatic in nature, and use a piston to force a fluid or gas through a restricted opening. In a fluidic system, the amount of dampening provided is substantially dependent upon the viscosity of the fluid. In a pneumatic system, the amount of dampening is substantially dependent upon the volume of gas being expelled. In both systems, the size of the restricted opening also determines the amount of dampening. In a typical spacecraft environment, the temperature differential can be as large as 200 degrees Celsius. Such a large temperature change can affect the viscosity of a fluid by a factor of 1000, and can have a substantial affect on gaseous volume and the effective area of the restricted opening. Because the damper must be designed so as not to overdamp the system under worst-case (e.g. maximum friction) conditions, the structure must be designed for the case of an underdamped system under opposite conditions (e.g. minimum friction). That is, a conventional spring-damper deployment system is over-designed to assure sufficient deployment force, then over-designed to minimize the effects caused by the over-designed deployment force. This combined over-design requires more massive hinges and stopping structures, again introducing additional design considerations and tradeoffs in a spacecraft design.